Combustion turbines comprise a casing or cylinder for housing a compressor section, combustion section and turbine section. The compressor section comprises an inlet end and a discharge end. The combustion section comprises an inlet end and a combustor transition. The combustor transition is proximate the discharge end of the combustion section and comprises a wall which defines a flow channel which directs the working gas into the turbine inlet end.
A supply of air is compressed in the compressor section and directed into the combustion section. The compressed air enters the combustion inlet and is mixed with fuel. The air/fuel mixture is then combusted to produce high temperature and high pressure gas. This working gas is then ejected past the combustor transition and injected into the turbine section to run the turbine.
The turbine section comprises rows of vanes which direct the working gas to the airfoil portions of the turbine blades. The working gas flows through the turbine section causing the turbine blades to rotate, thereby turning the rotor, which is connected to a generator for producing electricity.
As those skilled in the art are aware, the maximum power output of a gas turbine is achieved by heating the gas flowing through the combustion section to as high a temperature as is feasible. The hot gas, however, heats the various turbine components, such as the transition, vanes and ring segments, that it passes when flowing through the turbine.
Accordingly, the ability to increase the combustion firing temperature is limited by the ability of the turbine components to withstand increased temperatures. Consequently, various cooling methods have been developed to cool turbine hot parts. These methods include open-loop air cooling techniques and closed-loop cooling systems.
Conventional open-loop air cooling techniques divert air from the compressor to the combustor transition to cool the turbine hot parts. The cooling fluid extracts heat from the turbine components and then transfers into the inner transition flow channel and merges with the working fluid flowing into the turbine section. One drawback to open-loop cooling systems is that it diverts much needed air from the compressor, e.g., a significant amount of air flow is needed to keep the flame temperature of the combustor low. It is, therefore, desirable to provide a cooling system that diverts less air from the compressor.
Steam cooling of the vanes of stator blades is not new and has been the subject matter of commonly-assigned U.S. Pat. No. 5,320,483, of which a co-inventor is the inventor of the present invention. In combined cycle operation, steam at several pressure and temperature levels is readily available and it can be used to replace air as the cooling medium to cool the turbine hot parts.
The purpose of the present invention is to improve upon the present state of cooling of stator vanes of a gas turbine, particularly the first row vane. The operational requirements for such a design include a gas pressure range from 400 to 2000 psia, with a gas recovery temperature of approximately 2800.degree. F. operating in the transonic flow regime. The external gas path heat transfer coefficients assume a peak value of 1600 BTU at a point of highest curvature around the airfoil of the vane.
In addition to satisfying the above technical requirements for cooling a first row vane, the present invention is intended to (1) maintain simplicity for ease of casting the vane, (2) reduce the number of manufacturing operations, (3) reduce the number of parts, (4) be interchangeable with other advanced designs of different configuration, (5) use traditional cooling methods, and (6) achieve a minimum low cycle fatigue life. It is thus, desirable, to provide a versatile and effective first row vane design that lowers costs associated with manufacturing and maintenance.